The functions of Earth orbiting and other spacecraft require ever-increasing amounts of power as their functions and the complexity of those functions have evolved. Efforts are made to maximize the efficiency of energy use in performing these functions, but the amount of residual thermal energy required to be removed from the spacecraft structure in order to maintain a stable temperature has been increasing in designs made over a period of years. It is expected that the amount of residual power required to be removed from the spacecraft structure in the case of direct broadcast satellites will be greater than that required to be removed from current communications satellites.
Thermal energy cannot be removed from a spacecraft vehicle by conduction or convection, but only by radiation. Thus, the sources of unwanted heat on the spacecraft must be thermally coupled to radiators of sufficient size to maintain a satisfactory spacecraft temperature. In the context of a spacecraft, this thermal coupling presents unique problems. Because of the large cost of the spacecraft and of its launching, the various portions of the spacecraft which relate to the performance of its function must be highly reliable so that the cost may be amortized over the full design lifetime. Furthermore, the very large cost associated with launching the vehicle together with the desirability of maximizing the payload makes the weight of each structure of vital concern. Direct thermal coupling between the source of waste heat and its radiator is light in weight if the thermal path length is short, and is also extremely reliable. However, functional considerations may require a thermal path length which is so long that a direct thermally conductive path becomes heavier than other possible options. Heat transfer by the flow of fluid coolant between the source of waste heat and a heat radiating structure is often used. Because spacecraft travel through a flux of micrometeroids, there exists a danger that a pipe or channel through which coolant flows may be punctured, thereby resulting in the escape of coolant and loss of cooling capacity. This problem has been solved in the past by the use of a plurality of heat pipes thermally connecting the source of waste heat with the structure of the heat radiating element. Failure of one out of N heat pipes due to penetration by a micrometeroid causes a reduction in the capacity of the heat transfer system by a factor of 1/N, and does not result in total failure. The wall thickness of the heat pipes is selected by considerations of micrometeroid flux density and the desired reliability and life span.
Spacecraft having orbital paths inclined by less than about 30.degree. relative to the equatorial plane experience a lesser micrometeroid flux density than do spacecraft in polar or nearly-polar orbits. It is desirable to reduce the cost and weight of heat radiators.